Circumferential fuel shifting and biasing in an axial staged combustor for a gas turbine engine

ABSTRACT

An injector of a radial fuel injection system for a combustor of a gas turbine engine includes a swirler; and a fuel nozzle located within the swirler, the fuel nozzle located within the swirler, the fuel nozzle operable to provide a biased circumferential fuel distribution within the swirler. A method of controlling a fuel flow to a combustor of a gas turbine engine includes selectively controlling a biased circumferential fuel distribution within a swirler of a radial fuel injection system.

BACKGROUND

The present disclosure relates to a gas turbine engine and, moreparticularly, to a combustor section therefor.

Gas turbine engines, such as those which power modern commercial andmilitary aircrafts, include a compressor for pressurizing a supply ofair, a combustor for burning a hydrocarbon fuel in the presence of thepressurized air, and a turbine for extracting energy from the resultantcombustion gases. The combustor generally includes radially spaced apartinner and outer liners that define an annular combustion chambertherebetween. Arrays of circumferentially distributed combustion airholes penetrate multiple axial locations along each liner to radiallyadmit the pressurized air into the combustion chamber. A plurality ofcircumferentially distributed fuel injector orifices axially projectinto a forward section of the combustion chamber to supply the fuel formixing with the pressurized air.

Combustion of hydrocarbon fuel in the presence of pressurized air mayproduce nitrogen oxide (NOX) emissions that are subject to excessivelystringent controls by regulatory authorities, and thus may be sought tobe minimized. Lean-staged liquid-fueled aeroengine combustors canprovide low NO_(X) and particulate matter emissions, but are also proneto combustion instabilities. There are several mechanism that may causecombustion instabilities in radial-staged lean combustors including heatrelease concentrated in the front of the combustor, and weak flameholding at certain operating conditions where main stage air dilutes thepilot stage fuel-air ratio.

SUMMARY

An injector of a radial fuel injection system for a combustor of a gasturbine engine according to one disclosed non-limiting embodiment of thepresent disclosure can include a swirler and a fuel nozzle locatedwithin the swirler, the fuel nozzle operable to provide a biasedcircumferential fuel distribution within the swirler.

A further embodiment of the present disclosure may include, wherein thebiased circumferential fuel distribution within the swirler provides afuel bias to fuel enrich a circumferential location with respect to anexit diameter of the swirler.

A further embodiment of the present disclosure may include, wherein thebiased circumferential fuel distribution within the swirler provides afuel enriched bias to fuel enrich an outer portion of the jet withrespect to an exit diameter of the swirler.

A further embodiment of the present disclosure may include, wherein thebiased circumferential fuel distribution within the swirler provides afuel enriched bias to fuel enrich a jet leading edge with respect to across-flow of hot combustion products from an upstream pilot zone.

A further embodiment of the present disclosure may include, wherein thebiased circumferential fuel distribution within the swirler provides afuel enriched bias with respect to fuel enrich a wake with respect to across-flow of hot combustion products from an upstream pilot zone.

A further embodiment of the present disclosure may include, wherein thebiased circumferential fuel distribution within the swirler provides afuel enriched bias with respect to a flameholder.

A further embodiment of the present disclosure may include, wherein thefuel nozzle includes a first set of fuel injector orifices and a secondset of fuel injector orifices in the fuel nozzle, the second set of fuelinjector orifices downstream of the first set of fuel injector orificeswith respect to the swirler.

A further embodiment of the present disclosure may include a first fuelmanifold in communication with the first set of fuel injector orificesand a second fuel manifold in communication with the second set of fuelinjector orifices.

A further embodiment of the present disclosure may include a fueldivider valve that selectively communicates a respective percentage offuel flow to the first set of fuel injector orifices and the second setof fuel injector orifices to provide a selective flow split between thefirst set of fuel injector orifices and the second set of fuel injectororifices.

A further embodiment of the present disclosure may include, wherein thefuel divider valve is mechanical.

A further embodiment of the present disclosure may include, wherein thefuel divider valve is electrical.

An injector of a radial fuel injection system for a combustor of a gasturbine engine according to another disclosed non-limiting embodiment ofthe present disclosure can include a swirler; and a fuel nozzle locatedwithin the swirler, the fuel nozzle includes a first set of fuelinjector orifices and a second set of fuel injector orifices, the secondset of fuel injector orifices downstream of the first set of fuelinjector orifices with respect to the swirler, at least one of the setsof fuel injector orifices arranged to provide a biased circumferentialfuel distribution within the swirler.

A further embodiment of the present disclosure may include, wherein atleast one of the sets of fuel injector orifices is arranged to provide afuel enriched bias to fuel enrich a circumferential location withrespect to an exit diameter of the swirler.

A further embodiment of the present disclosure may include, wherein atleast one of the sets of fuel injector orifices is arranged to provide afuel enriched bias to fuel enrich an outer portion of the jet withrespect to an exit diameter of the swirler.

A further embodiment of the present disclosure may include, wherein atleast one of the sets of fuel injector orifices is arranged to provide afuel enriched bias to fuel enrich a jet leading edge from the swirlerwith respect to a cross-flow of hot combustion products from an upstreampilot zone.

A further embodiment of the present disclosure may include, wherein atleast one of the sets of fuel injector orifices is arranged to provide afuel enriched bias with respect to fuel enrich a wake from the swirlerwith respect to a cross-flow of hot combustion products from an upstreampilot zone.

A further embodiment of the present disclosure may include, wherein atleast one of the sets of fuel injector orifices is arranged to provide afuel enriched bias with respect to a flameholder.

A further embodiment of the present disclosure may include, furthercomprising a first fuel manifold in communication with the first set offuel injector orifices and a second fuel manifold in communication withthe second set of fuel injector orifices.

A further embodiment of the present disclosure may include a fueldivider valve that selectively communicates a respective percentage offuel flow to the first set of fuel injector orifices and the second setof fuel injector orifices to provide a selective flow split between thefirst set of fuel injector orifices and the second set of fuel injectororifices.

A further embodiment of the present disclosure may include, wherein thefuel divider valve is at least one of a mechanical or electrical dividervalve.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiment. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a schematic cross-section of an example gas turbine enginearchitecture;

FIG. 2 is an expanded longitudinal schematic sectional view of acombustor section according to one non-limiting embodiment that may beused with the example gas turbine engine

FIG. 3 is a perspective partial longitudinal sectional view of thecombustor section;

FIG. 4 is a schematic longitudinal sectional view of the combustorsection which illustrates a forward axial fuel injection system and adownstream radial fuel injections system according to one disclosednon-limiting embodiment;

FIG. 5 is a schematic lateral sectional view of a combustor whichillustrates an in-line fuel nozzle arrangement according to anotherdisclosed non-limiting embodiment;

FIG. 6 is a schematic lateral sectional view of a combustor whichillustrates a clocked fuel nozzle arrangement according to anotherdisclosed non-limiting embodiment;

FIG. 7 is a schematic longitudinal sectional view of a combustor whichillustrates a tangential fuel nozzle arrangement according to anotherdisclosed non-limiting embodiment;

FIG. 8 is a schematic lateral sectional view of a combustor whichillustrates a tangential fuel nozzle arrangement according to anotherdisclosed non-limiting embodiment;

FIG. 9 is a schematic longitudinal sectional view of a combustor whichillustrates an axially angled fuel nozzle arrangement according toanother disclosed non-limiting embodiment;

FIG. 10 is a schematic longitudinal sectional view of a combustor whichillustrates an outer radial fuel injection system arrangement accordingto another disclosed non-limiting embodiment;

FIG. 11 is a schematic longitudinal sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment;

FIG. 12 is a schematic longitudinal sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment;

FIG. 13 is a schematic longitudinal sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment;

FIG. 14 is a schematic lateral sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment;

FIG. 15 is a schematic lateral sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment;

FIG. 16 is a schematic lateral sectional view of a combustor whichillustrates the axial and radial fuel injection systems that include anumerically different relationship according to another disclosednon-limiting embodiment;

FIG. 17 is a schematic longitudinal sectional view of a combustor whichillustrates a relationship between the axial and radial fuel injectionsystems according to another disclosed non-limiting embodiment; and

FIG. 18 is a schematic view of a fuel injector of a radial fuelinjection system according to one disclosed non-limiting embodiment;

FIG. 19 is a schematic view of a fuel manifold for a first set of fuelinjector orifices and a second set of fuel injector orifices;

FIG. 20 is a graphical representation of a fuel flow split between afirst set of fuel injector orifices and a second set of fuel injectororifices;

FIG. 21 is a graphical representation of a radial fuel distribution;

FIG. 22 is a graphical representation of a radial F/A distribution;

FIG. 23 is a schematic view of a fuel injector of a radial fuelinjection system with a flameholder according to one disclosednon-limiting embodiment;

FIG. 24 is a schematic view of a fuel orifice arrangement according toone disclosed non-limiting embodiment;

FIG. 25 is a schematic view of a fuel orifice arrangement according toone disclosed non-limiting embodiment;

FIG. 26 is a schematic view of a fuel orifice arrangement according toone disclosed non-limiting embodiment;

FIG. 27 is a schematic view of a fuel orifice arrangement according toone disclosed non-limiting embodiment;

FIG. 28 is a schematic view of a fuel orifice arrangement according toone disclosed non-limiting embodiment;

FIG. 29 is a graphical representation of an example circumferentiallyfuel distribution; and

FIG. 30 is a graphical representation of an example circumferentiallyF/A distribution.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbo fan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flowpath while the compressor section 24 drives airalong a core flowpath for compression and communication into thecombustor section 26 then expansion through the turbine section 28.Although depicted as a turbofan in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with turbofans as the teachings may be applied toother types of turbine engines such as a turbojets, turboshafts, andthree-spool (plus fan) turbofans wherein an intermediate spool includesan intermediate pressure compressor (“IPC”) between a Low PressureCompressor (“LPC”) and a High Pressure Compressor (“HPC”), and anintermediate pressure turbine (“IPT”) between the high pressure turbine(“HPT”) and the Low pressure Turbine (“LPT”).

The engine 20 generally includes a low spool 30 and a high spool 32mounted for rotation about an engine central longitudinal axis Arelative to an engine static structure 36 via several bearing structures38. The low spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor (“LPC”) 44 and a lowpressure turbine (“LPT”) 46. The inner shaft 40 drives the fan 42directly or through a geared architecture 48 to drive the fan 42 at alower speed than the low spool 30. An exemplary reduction transmissionis an epicyclic transmission, namely a planetary or star gear system.

The high spool 32 includes an outer shaft 50 that interconnects a highpressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54. Acombustor 56 is arranged between the high pressure compressor 52 and thehigh pressure turbine 54. The inner shaft 40 and the outer shaft 50 areconcentric and rotate about the engine central longitudinal axis A whichis collinear with their longitudinal axes.

Core airflow is compressed by the LPC 44 then the HPC 52, mixed with thefuel and burned in the combustor 56, then expanded over the HPT 54 andthe LPT 46. The turbines 54, 46 rotationally drive the respective lowspool 30 and high spool 32 in response to the expansion. The main engineshafts 40, 50 are supported at a plurality of points by bearingstructures 38 within the static structure 36. It should be understoodthat various bearing structures 38 at various locations mayalternatively or additionally be provided.

In one non-limiting example, the gas turbine engine 20 is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20bypass ratio is greater than about six (6:1). The geared architecture 48can include an epicyclic gear train, such as a planetary gear system orother gear system. The example epicyclic gear train has a gear reductionratio of greater than about 2.3, and in another example is greater thanabout 2.5:1. The geared turbofan enables operation of the low spool 30at higher speeds which can increase the operational efficiency of theLPC 44 and LPT 46 and render increased pressure in a fewer number ofstages.

A pressure ratio associated with the LPT 46 is pressure measured priorto the inlet of the LPT 46 as related to the pressure at the outlet ofthe LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. Inone non-limiting embodiment, the bypass ratio of the gas turbine engine20 is greater than about ten (10:1), the fan diameter is significantlylarger than that of the LPC 44, and the LPT 46 has a pressure ratio thatis greater than about five (5:1). It should be understood, however, thatthe above parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by thebypass flow path due to the high bypass ratio. The fan section 22 of thegas turbine engine 20 is designed for a particular flightcondition—typically cruise at about 0.8 Mach and about 35,000 feet(10668 m). This flight condition, with the gas turbine engine 20 at itsbest fuel consumption, is also known as bucket cruise Thrust SpecificFuel Consumption (TSFC). TSFC is an industry standard parameter of fuelconsumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed isthe actual fan tip speed divided by an industry standard temperaturecorrection of (“Tram”/518.7)^(0.5). The Low Corrected Fan Tip Speedaccording to one non-limiting embodiment of the example gas turbineengine 20 is less than about 1150 fps (351 m/s).

With reference to FIG. 2, the combustor section 26 generally includes acombustor 56 with an outer combustor liner assembly 60, an innercombustor liner assembly 62 and a diffuser case module 64. The outercombustor liner assembly 60 and the inner combustor liner assembly 62are spaced apart such that a combustion chamber 66 is definedtherebetween. The combustion chamber 66 is generally annular in shape.

The outer combustor liner assembly 60 is spaced radially inward from anouter diffuser case 64-O of the diffuser case module 64 to define anouter annular plenum 76. The inner combustor liner assembly 62 is spacedradially outward from an inner diffuser case 64-I of the diffuser casemodule 64 to define an inner annular plenum 78. It should be understoodthat although a particular combustor is illustrated, other combustortypes with various combustor liner arrangements will also benefitherefrom. It should be further understood that the disclosed coolingflow paths are but an illustrated embodiment and should not be limitedonly thereto.

The combustor liner assemblies 60, 62 contain the combustion productsfor direction toward the turbine section 28. Each combustor linerassembly 60, 62 generally includes a respective support shell 68, 70which supports one or more liner panels 72, 74 mounted to a hot side ofthe respective support shell 68, 70. Each of the liner panels 72, 74 maybe generally rectilinear and manufactured of, for example, a nickelbased super alloy, ceramic or other temperature resistant material andare arranged to form a liner array. In one disclosed non-limitingembodiment, the liner array includes a multiple of forward liner panels72A and a multiple of aft liner panels 72B that are circumferentiallystaggered to line the hot side of the outer shell 68 (also shown in FIG.4). A multiple of forward liner panels 74A and a multiple of aft linerpanels 74B are circumferentially staggered to line the hot side of theinner shell 70 (also shown in FIG. 3).

The combustor 56 further includes a forward assembly 80 immediatelydownstream of the compressor section 24 to receive compressed airflowtherefrom. The forward assembly 80 generally includes an annular hood82, a bulkhead assembly 84, a multiple of forward fuel nozzles 86 (oneshown) and a multiple of swirlers 90 (one shown). The multiple of fuelnozzles 86 (one shown) and the multiple of swirlers 90 (one shown)define an axial pilot fuel injection system 92 that directs the fuel-airmixture into the combustor chamber generally along an axis F.

The bulkhead assembly 84 includes a bulkhead support shell 96 secured tothe combustor liner assemblies 60, 62, and a multiple ofcircumferentially distributed bulkhead liner panels 98 secured to thebulkhead support shell 96. The annular hood 82 extends radially between,and is secured to, the forwardmost ends of the combustor linerassemblies 60, 62. The annular hood 82 includes a multiple ofcircumferentially distributed hood ports 94 that accommodate therespective forward fuel nozzles 86 and direct air into the forward endof the combustion chamber 66 through a respective swirler 90. Eachforward fuel nozzle 86 may be secured to the diffuser case module 64 andproject through one of the hood ports 94 and through the respectiveswirler 90. Each of the fuel nozzles 86 is directed through therespective swirler 90 and the bulkhead assembly 84 along a respectiveaxis F.

The forward assembly 80 introduces core combustion air into the forwardsection of the combustion chamber 66 while the remainder enters theouter annular plenum 76 and the inner annular plenum 78. The multiple offuel nozzles 86 and adjacent structure generate a blended fuel-airmixture that supports stable combustion in the combustion chamber 66.

Opposite the forward assembly 80, the outer and inner support shells 68,70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in theHPT 54 to define a combustor exit 100. The NGVs 54A are static enginecomponents which direct core airflow combustion gases onto the turbineblades of the first turbine rotor in the turbine section 28 tofacilitate the conversion of pressure energy into kinetic energy. Thecombustion gases are also accelerated by the NGVs 54A because of theirconvergent shape and are typically given a “spin” or a “swirl” in thedirection of turbine rotor rotation. The turbine rotor blades absorbthis energy to drive the turbine rotor at high speed.

With reference to FIG. 3, a multiple of cooling impingement holes 104penetrate through the support shells 68, 70 to allow air from therespective annular plenums 76, 78 to enter cavities 106A, 106B formed inthe combustor liner assemblies 60, 62 between the respective supportshells 68, 70 and liner panels 72, 74. The cooling impingement holes 104are generally normal to the surface of the liner panels 72, 74. The airin the cavities 106A, 106B provides cold side impingement cooling of theliner panels 72, 74 that is generally defined herein as heat removal viainternal convection.

A multiple of cooling film holes 108 penetrate through each of the linerpanels 72, 74. The geometry of the film holes, e.g., diameter, shape,density, surface angle, incidence angle, etc., as well as the locationof the holes with respect to the high temperature main flow alsocontributes to effusion film cooling. The liner panels 72, 74 with acombination of impingement holes 104 and film holes 108 may sometimes bereferred to as an Impingement Film Floatliner assembly. It should beappreciated that other liner panel assemblies inclusive of a singlepanel.

The cooling film holes 108 allow the air to pass from the cavities 106A,106B defined in part by a cold side 110 of the liner panels 72, 74 to ahot side 112 of the liner panels 72, 74 and thereby facilitate theformation of a film of cooling air along the hot side 112. The coolingfilm holes 108 are generally more numerous than the impingement holes104 to promote the development of a film cooling along the hot side 112to sheath the liner panels 72, 74. Film cooling as defined herein is theintroduction of a relatively cooler airflow at one or more discretelocations along a surface exposed to a high temperature environment toprotect that surface in the immediate region of the airflow injection aswell as downstream thereof.

A multiple of dilution holes 116 may penetrate through both therespective support shells 68, 70 and liner panels 72, 74 along a commonaxis downstream of the forward assembly 80 to quench the hot gases bysupplying cooling air radially into the combustor. That is, the multipleof dilution holes 116 provide a direct path for airflow from the annularplenums 76, 78 into the combustion chamber 66.

With reference to FIG. 4, a radial main fuel injection system 120communicates with the combustion chamber 66 downstream of the axialpilot fuel injection system 92 generally transverse to axis F of anAxially Controlled Stoichiometry (ACS) Combustor. The radial main fuelinjection system 120 introduces a portion of the fuel required fordesired combustion performance, e.g., emissions, operability,durability, as well as to lean-out the fuel contribution provided by theaxial pilot fuel injection system 92. In one disclosed non-limitingembodiment, the radial main fuel injection system 120 is positioneddownstream of the axial pilot fuel injection system 92 and upstream ofthe multiple of dilution holes 116.

The radial main fuel injection system 120 generally includes a radiallyouter fuel injection manifold 122 (illustrated schematically) and/or aradially inner fuel injection manifold 124 (illustrated schematically)with a respective multiple of outer fuel nozzles 126 and a multiple ofinner fuel nozzles 128. The radially outer fuel injection manifold 122and/or a radially inner fuel injection manifold 124 may be mounted tothe diffuser case module 64 and/or to the shell 68, 70, however, variousmount arrangements may alternatively or additionally provided.

Each of the multiple of outer and inner fuel nozzles 126, 128 arelocated within a respective mixer 130, 132 to mix the supply of fuelinto the pressurized air within the diffuser case module 64. As definedherein, a “mixer” as compared to a “swirler” may generate, for example,zero swirl, a counter-rotating swirl, a specific swirl which provides aresultant swirl or a residual net swirl which may be further directed atan angle. It should be appreciated that various combinations thereof mayalternatively be utilized.

The radial main fuel injection system 120 may include only the radiallyouter fuel injection manifold 96 with the multiple of outer fuel nozzles126; only the radially inner fuel injection manifold 124 with themultiple of inner fuel nozzles 128; or both (shown). It should beappreciated that the radial main fuel injection system 120 may includesingle sets of outer fuel nozzles 126 and inner fuel nozzles 128 (shown)or multiple axially distributed sets of, for example, relatively smallerfuel nozzles.

The radial main fuel injection system 120 may be circumferentiallyarranged in a multiple of configurations. In one disclosed non-limitingembodiment, the multiple of outer fuel nozzles 126 and the multiple ofinner fuel nozzles 128 are circumferentially arranged so that thenozzles 126, 128 are directly opposed (FIG. 6). In another disclosednon-limiting embodiment, the multiple of outer fuel nozzles 126 and themultiple of inner fuel nozzles 128 are circumferentially staggered sothat the nozzles 126, 128 are not directly opposed (FIG. 6).Furthermore, the nozzles 126, 128 may be angled perpendicularly (FIG.7), tangentially (FIG. 8), or at an angle such as downstream (FIG. 9)relative to the cross flow from the fuel nozzles 86 of the axial pilotfuel injection system 92 that are directed along axis F.

Alternatively still, the multiple of outer fuel nozzles 126 may bepositioned through the outer liner 72 opposite or staggered relative toa non-fueled mixer 132′ on the inner liner 74 (FIG. 10). That is, thenon-fueled mixer 132′ provides airflow but not fuel.

The lean-staging is accomplished by axially distributing the fuelinjection with a front-end pilot injector and a downstream main injectorto axially distribute the heat release similar to an RQL designs, butwith lean/lean combustion to enable low NOx and PM emissions. This isdifferent than radial staged designs where all the fuel is injected atthe front-end of the combustor. Moving the heat release away from thefront-end can be a pressure anti-node for longitudinal acoustic modes todecrease coupling with these modes.

With respect to FIG. 12, the forward fuel nozzles 86 arecircumferentially spaced apart between about 80%-200% of a bulkheadheight B. The bulkhead height B as defined herein is the radial distancebetween the liner panels 72, 74 at the forward end of the combustionchamber 66 at the bulkhead liner panels 98 of bulkhead assembly 84. Themultiple of outer fuel nozzles 126 and the inner fuel nozzles 128 areaxially spaced a distance D between 50%-150% of the bulkhead height Baft of the forward fuel nozzles 86.

The multiple of outer fuel nozzles 126 are radially spaced a distance Rfrom the inner fuel nozzles 128 at between about 100%-200% of thebulkhead height B. It should be understood that the distance R may bewith respect to the liner panels 72, 74 should the radial main fuelinjection system 120 only utilize outer fuel nozzles 126 (FIG. 12) orinner fuel nozzles 128 (FIG. 13).

With respect to FIG. 14, the multiple of outer fuel nozzles 126 andmultiple of inner fuel nozzles 128 may be arranged circumferentiallyin-line with the forward fuel nozzles 86. Alternatively, the multiple ofouter fuel nozzles 126 and multiple of inner fuel nozzles 128 may bearranged circumferentially between the forward fuel nozzles 86 at, forexample, quarter pitch (FIG. 15). The multiple of outer fuel nozzles 126and/or the multiple of inner fuel nozzles 128 may be spaced apart adistance C of between 25%-100% of the bulkhead height Bcircumferentially, which alternatively, may be defined as about 1.5-5fuel jet diameters. It should be appreciated that variouscircumferential and other relationships may be utilized and that fueljet diameter and bulkhead sizing are but examples thereof.

Alternatively still, with respect to FIG. 16, the multiple of outer fuelnozzles 126 may be more numerous than the forward fuel nozzles 86. Inthis disclosed non-limiting embodiment, twice the number of outer fuelnozzles 126 as compared to the forward fuel nozzles 86. The multiple ofouter fuel nozzles 126 include both in-line and circumferentiallydistributed forward fuel nozzles 86

With reference to FIG. 17, the axial pilot fuel injection system 92, theradial main fuel injection system 120 and the multiple of dilution holes116 define a forward combustion zone 140 axially between the bulkheadassembly 84 and the forward section of the radial main fuel injectionsystem 120, as well as a downstream combustion zone 142 between theforward section of the radial main fuel injection system 120 and thecombustor exit 100. The downstream combustion zone 142 is axiallyproximate the multiple of dilution holes 116.

In one disclosed non-limiting embodiment, the axial pilot fuel injectionsystem 92 provides about 10%-35% of the combustor airflow, the radialmain fuel injection system 120 provides about 30%-60% of combustorairflow while the multiple of dilution holes 116 provide about 5%-20% ofthe combustor airflow. It should be appreciated that these ranges ofcombustor airflow may define a total combustor airflow less than 100%with the remainder being cooling airflow. It should be furtherappreciated that generally as the combustor airflow from the axial pilotfuel injection system 92 increases, the radial main fuel injectionsystem 120 decreases and vice-versa with the balance being from themultiple of dilution holes 116. In one specific example, the axial pilotfuel injection system 92 provides about 20% of the combustor airflow,the radial main fuel injection system 120 provides about 45% ofcombustor airflow while the multiple of dilution holes 116 provide about10% of the combustor airflow with the remainder being cooling airflow.

In one disclosed non-limiting embodiment, the forward combustion zone140 defines about 20% to 50% of the total combustor chamber 66 volumeand the downstream combustion zone 142 defines about 50% to 80% of thetotal combustor chamber 66 volume.

In one disclosed non-limiting embodiment, the downstream combustion zone142 forms an axial length L of about 100%-250% a height H of thecombustion chamber 66 between the liners 72, 74 at the radial main fuelinjection system 120 location. The height H as defined herein is theradial distance between the liner panels 72, 74 within the combustionchamber 66 proximate the radial main fuel injection system 120 location.It should be appreciated that various combinations of theabove-described geometries may be provided.

With reference to FIG. 18, one example injector 200 of the radial mainfuel injection system 120 includes a fuel nozzle 202 located within arespective swirler 204. It should be appreciated that although oneexample injector 200 of the radial main fuel injection system 120 isdepicted various other injectors may benefit herefrom.

The fuel nozzle 202 includes a multiple of fuel injector orifices 206that introduce a premixed or partially premixed fuel/air jet J into across-flow F of hot combustion products created in the upstream pilotregion of the combustion chamber 66. The swirler 204 in this embodimentis an axial swirler with a central fuel nozzle 202.

In this embodiment, the multiple of fuel injector orifices 206 include afirst set of fuel injector orifices 206A and a second set of fuelinjector orifices 206B that are located downstream of the first set offuel injector orifices 206A. The fuel injector orifices 206 are locateddownstream of the swirler 204 with respect to the combustion chamber 66.The first set of fuel injector orifices 206A and the second set of fuelinjector orifices 206B each receive fuel independently from a respectivefirst fuel manifold 208A and second fuel manifold 208B. A fuel dividervalve 210 selectively communicates a respective percentage of fuel flowto the first set of fuel injector orifices 206A and the second set offuel injector orifices 206B (FIG. 20). The fuel divider valve 210 may bea mechanical or electrical flow divider valve to shift the flow betweenmanifolds 208A, 208B to provide a selective flow split between the firstset of fuel injector orifices 206A, and the second set of fuel injectororifices 206B (FIG. 21).

With reference to FIGS. 21 and 22, a mixer with a 50/50 design pointfuel flow split provides a fuel jet penetration at 55% (0.55 r/R) ofswirler radius. Fuel shifting between the first set of fuel injectororifices 206A and the second set of fuel injector orifices 206B can beuse to tailor the radial F/A distribution profile at the exit of theswirler 204 in order to anchor the flame (FIG. 23). Shifting fuel splitto 75/25 provides a fuel jet penetration of the first set of fuelinjector orifices 206A of 0.275 r/R and 0.825 r/R for the second set offuel injector orifices 206B to move the fuel jet radially outward whichresults in a shifted radial F/A distribution toward the outer diameter(FIG. 22).

With reference to FIG. 23, in another disclosed non-limiting embodiment,a flameholder 220, such as a V-gutter may be located at the exit of theswirler 204. The flameholder 220 may be aligned with the cross-flow F ofthe hot combustion products from the upstream pilot zone through thecombustion chamber 66.

With reference to FIGS. 24, in another disclosed non-limitingembodiment, clocking of the first set of fuel injector orifices 206A andthe second set of fuel injector orifices 206B may be staggered by ½ theangular hole spacing. That is, the first set of fuel injector orifices206A and the second set of fuel injector orifices 206B need not be indifferent axial locations and may be coplanar at the fuel nozzle 202 ata constant axial distance from the swirler 204.

With reference to FIGS. 25, in another disclosed non-limitingembodiment, one or more, or each, of the first set of fuel injectororifices 206A, and the second set of fuel injector orifices 206B may beclocked to compensate for the swirl within the swirler 204. That is, thefuel injection location is clocked at an injection direction such thatthe fuel-enriched region is rotated by the swirl generated by theswirler 204 and the partially premixed fuel/air jet J directed into thecross-flow F of hot combustion products is clocked to a desiredresultant location at the swirler exit. The injection direction isrotated as required to account for swirl such as, for example, by 80degree clockwise to account for 80 degree counter-clockwise rotation ofthe flow from the injector to the swirler exit. The fuel injectororifices may be clocked to, for example, enrich the jet leading edge(FIG. 25), or to enrich the jet wake (FIG. 26). The fuel injectororifices may alternatively be clocked to be aligned with the flameholder220 to enrich the flameholder wake (FIG. 27).

With reference to FIGS. 28, in another disclosed non-limitingembodiment, the first set of fuel injector orifices 206A and the secondset of fuel injector orifices 206B receive fuel independently from therespective first fuel manifold 208A and the second fuel manifold 208B(FIG. 19). That is, of six (6) total fuel injector orifices, the firstset of fuel injector orifices 206A may include five (5) injectors thateach receive a first fuel flow while the second set of fuel injectororifices 206B may include one (1) injector that receives a second fuelflow to provide a fuel bias to enrich particular circumferentiallocations of the reacting jet in the cross-flow at a desiredcircumferential location. In this embodiment, the sets of fuel injectororifices are defined by a particular number of orifices at a commonradial position with respect to the swirler. The particular fuel biasinginjectors do not have to be in a separate axial location within swirler204 and can be co-located axially as one or two of the typically sixinjectors that are arranged circumferentially.

With reference to FIGS. 29 and 30, an example nominal uniform fueldistribution is provided to six (6) fuel injection orifices, each with⅙=17% of total flow. For example, by biasing fuel flow as follows: 79%to five (5) injectors (16% each) and 21% to one (1) injector (16+5=21%;a 5% bias) an increased fuel flow to one injector by +5% increases theF/A at a desired circumferential location. That is, the Fuel/Air Ratiois significantly increased at targeted circumferential location.

The injector 200 provides for stable and robust anchoring/flameholdingof the main zone reacting jet, which facilitates mitigation ofcombustion dynamics, improved dynamic stability, and prevention ofintermittent flame lift-off. Local fuel enrichment at the outer diameterof the swirler 204 exit enhances flameholding. Local fuel enrichment atleading edge, and/or in the wake of jet, and/or aligned with theflameholder at the swirler exit will enhance flameholding. Fuel shiftingor fuel biasing can be used to create a richer F/A mixture at a specificlocation where the flame anchoring is desired. Fuel shifting and fuelbiasing for a liquid-fueled aero engine axially-staged lean-leancombustor configuration may be provided by radial fuel re-distributionwithin the swirler, and/or non-uniform circumferential distributionwithin or with respect to the swirler.

The use of the terms “a” and “an” and “the” and similar references inthe context of description (especially in the context of the followingclaims) are to be construed to cover both the singular and the plural,unless otherwise indicated herein or specifically contradicted bycontext. The modifier “about” used in connection with a quantity isinclusive of the stated value and has the meaning dictated by thecontext (e.g., it includes the degree of error associated withmeasurement of the particular quantity). All ranges disclosed herein areinclusive of the endpoints, and the endpoints are independentlycombinable with each other. It should be appreciated that relativepositional terms such as “forward,” “aft,” “upper,” “lower,” “above,”“below,” and the like are with reference to the normal operationalattitude of the vehicle and should not be considered otherwise limiting.

Although the different non-limiting embodiments have specificillustrated components, the embodiments of this invention are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be appreciated that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be appreciated that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beappreciated that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. An injector of a radial fuel injection system fora combustor of a gas turbine engine comprising: a swirler; and a fuelnozzle located within the swirler, the fuel nozzle operable to provide abiased circumferential fuel distribution within the swirler.
 2. Theinjector as recited in claim 1, wherein the biased circumferential fueldistribution within the swirler provides a fuel bias to fuel enrich acircumferential location with respect to an exit diameter of theswirler.
 3. The injector as recited in claim 1, wherein the biasedcircumferential fuel distribution within the swirler provides a fuelenriched bias to fuel enrich an outer portion of the jet with respect toan exit diameter of the swirler.
 4. The injector as recited in claim 1,wherein the biased circumferential fuel distribution within the swirlerprovides a fuel enriched bias to fuel enrich a jet leading edge withrespect to a cross-flow of hot combustion products from an upstreampilot zone.
 5. The injector as recited in claim 1, wherein the biasedcircumferential fuel distribution within the swirler provides a fuelenriched bias with respect to fuel enrich a wake with respect to across-flow of hot combustion products from an upstream pilot zone. 6.The injector as recited in claim 1, wherein the biased circumferentialfuel distribution within the swirler provides a fuel enriched bias withrespect to a flameholder.
 7. The injector as recited in claim 1, whereinthe fuel nozzle includes a first set of fuel injector orifices and asecond set of fuel injector orifices in the fuel nozzle, the second setof fuel injector orifices downstream of the first set of fuel injectororifices with respect to the swirler.
 8. The injector as recited inclaim 7, further comprising a first fuel manifold in communication withthe first set of fuel injector orifices and a second fuel manifold incommunication with the second set of fuel injector orifices.
 9. Theinjector as recited in claim 8, further comprising a fuel divider valvethat selectively communicates a respective percentage of fuel flow tothe first set of fuel injector orifices and the second set of fuelinjector orifices to provide a selective flow split between the firstset of fuel injector orifices and the second set of fuel injectororifices.
 10. The injector as recited in claim 9, wherein the fueldivider valve is mechanical.
 11. The injector as recited in claim 9,wherein the fuel divider valve is electrical.
 12. An injector of aradial fuel injection system for a combustor of a gas turbine enginecomprising: a swirler; and a fuel nozzle located within the swirler, thefuel nozzle includes a first set of fuel injector orifices and a secondset of fuel injector orifices, the second set of fuel injector orificesdownstream of the first set of fuel injector orifices with respect tothe swirler, at least one of the sets of fuel injector orifices arrangedto provide a biased circumferential fuel distribution within theswirler.
 13. The injector as recited in claim 12, wherein at least oneof the sets of fuel injector orifices is arranged to provide a fuelenriched bias to fuel enrich a circumferential location with respect toan exit diameter of the swirler.
 14. The injector as recited in claim12, wherein at least one of the sets of fuel injector orifices isarranged to provide a fuel enriched bias to fuel enrich an outer portionof the jet with respect to an exit diameter of the swirler.
 15. Theinjector as recited in claim 12, wherein at least one of the sets offuel injector orifices is arranged to provide a fuel enriched bias tofuel enrich a jet leading edge from the swirler with respect to across-flow of hot combustion products from an upstream pilot zone. 16.The injector as recited in claim 12, wherein at least one of the sets offuel injector orifices is arranged to provide a fuel enriched bias withrespect to fuel enrich a wake from the swirler with respect to across-flow of hot combustion products from an upstream pilot zone. 17.The injector as recited in claim 12, wherein at least one of the sets offuel injector orifices is arranged to provide a fuel enriched bias withrespect to a flameholder.
 18. The injector as recited in claim 12,further comprising a first fuel manifold in communication with the firstset of fuel injector orifices and a second fuel manifold incommunication with the second set of fuel injector orifices.
 19. Theinjector as recited in claim 18, further comprising a fuel divider valvethat selectively communicates a respective percentage of fuel flow tothe first set of fuel injector orifices and the second set of fuelinjector orifices to provide a selective flow split between the firstset of fuel injector orifices and the second set of fuel injectororifices.
 20. The injector as recited in claim 19, wherein the fueldivider valve is at least one of a mechanical or electrical dividervalve.